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Role: Analyst, Scope Owner
Duration: September 2024 – August 2025
Objective: Develop a numerical model in Python for hybrid rocket engine performance, with the goal of generating a .RSE file to be used in Open Rocket for designing our sounding rocket.
Summary: To design the LOX Paraffin rocket engine for our 2024-2025 rocket I wrote a custom Python package. It utilizes NASA Chemical Equilibrium Analysis (CEA) in MATLAB to calculate combustion conditions during each time step in the engine based on a shifting O/F, shifting chamber pressure, and a combustion efficiency knockdown. It then applies compressible gas physics to calculate the thrust and specific impulse at a given altitude (or ambient pressure).
The script also calculates the center of gravity of the propulsion system, which is critical for ensuring in-flight stability. It is fully parameterized, allowing for hundreds of iterations to finely tune tank lengths, chamber lengths, and propellant mass with a given pressurant tank, fluids stack, valve, injector, and nozzle. The model is anchored by how we build our engines and the geometry of an amateur hybrid rocket motor.
It also generates a fully parameterized .RSE file, which is used for simulating the entire vehicle in Open Rocket. RSE files contain the thrust, mass, and CG for a given time step in the propulsion system. After performing a simulated flight in Open Rocket the air pressure vs. time can be used to re-run a .RSE, giving a slightly different thrust vs. time due to over or under expansion in the nozzle. The expansion ratio can then be modified to find the best solution, and an optimal set of parameters is found.
An initial model for tank pressurization was also integrated into the model to parameterize the COPV, regulators, and tubing. It used relations for choked and unchoked compressible flow to size the dome-loaded regulator. Test data showed that the model under-predicted ullage collapse, and additional modeling anchored with sensor data is underway.
Results: The engine model I developed was used to design the 2024-2025 motor. The first static fire yielded 1450 lbf of thrust, likely due to an overly constrictive injector and poor mixing when the baffle fractured. The model was adapted to generate RSE files for flight, and we achieved within 1000ft of our modeled apogee.
Final Parameters:
*** MASS PROPERTIES ***
Wet Mass: 48.69 kg
Dry Mass: 30.4245 kg
Dry CG: 1.2624 m
Wet CG: 1.33 m
Overall Length: 2.9172 m
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*** OXIDIZER TANK MASS PROPERTIES ***
Ox Tank Mass: 5.5221 kg
Ox Tank Length: 0.9208 m
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*** COMBUSTION CHAMBER MASS PROPERTIES ***
CC Shell Mass: 3.9267 kg
CC Length: 0.5745 m
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*** INITIAL CONDITIONS ***
Initial Oxidizer Mass: 12.4 kg
Initial Fuel Mass: 4.84 kg
Initial Pressurant Mass: 1.0255 kg
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*** FINAL CONDITIONS ***
Remaining Oxidizer: 0.0931 kg
Remaining Fuel: -0.0001 kg
Final Port Diameter: 0.1334 m
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*** OX CONDITIONS ***
Ox Tank Length: 0.9208 m
Ox Mass Flow Rate: 2.53 kg/s
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*** CHAMBER CONDITIONS ***
Grain Length: 0.5315 m
Initial Port Diameter: 0.07 m
Max Port Mass Flux: 657.4074 kg/m^2/s
Chamber Temperature: 3558.2446 K
Throat Temperature: 3342.0457 K
Exit Temperature: 2400.9292 K
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*** NOZZLE CONDITIONS ***
Throat Diameter: 50.9385 mm
Exit Diameter: 120.2193 mm
Average Throat Mass Flux: 1756.6287 kg/m^2/s
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*** PERFORMANCE ***
Average Thrust: 8523.0454 N
Max Thrust (N): 8679.7918 N
Max Thrust (lbf): 1951.2953 lbf
Burn Time: 4.789 s
Average Thrust Coefficient: 1.4967
Total Impulse: 40816.8087 Ns
Average ISP: 242.7332 s
See the LOX feed system page for additional footage of static fire and launch.